Gas turbine combustion chamber

ABSTRACT

The present invention relates to a gas-turbine combustion chamber with an outer combustion chamber wall concentric to a gas-turbine center axis and with an inner combustion chamber wall, with several burners arranged spread over the circumference of the combustion chamber, and with air inlet recesses which in at least one radial plane are provided spread over the circumference on an outer combustion chamber wall and on an inner combustion chamber wall, where the burner is designed to form a flow provided with a swirl and where air inlet recesses assigned to a burner are dimensioned in differing size in order to generate airflows of differing size, characterized in that at least one of the respective air inlet recesses is designed for supplying air in the flow direction of the swirl of the combustion chamber flow and is provided with flow guidance walls.

This invention relates to a gas-turbine combustion chamber in accordance with the features of the generic part of claim 1.

In detail, the invention relates to a gas-turbine combustion chamber with an outer combustion chamber wall concentric to a gas-turbine center axis and with an inner combustion chamber wall, with several burners arranged spread over the circumference of the combustion chamber, and with air inlet recesses which in at least one radial plane are provided spread over the circumference on the outer combustion chamber wall and on the inner combustion chamber wall, where the burner is designed to form a flow provided with a swirl and where air inlet recesses assigned to a burner are dimensioned in differing size in order to generate airflows of differing size.

For aircraft gas turbines, it has proved particularly important that they have low emission values, in particular with regard to NOx, CO, UHC and soot. In addition, the further requirements must be met, for example relating to the operating characteristics of the combustion chamber, such as ignitability and flame stability.

The state of the art shows designs in which first rich combustion and then lean combustion takes place (rich-burn-quick-quench-lean-burn (RQL)). This burner design used in modern aircraft gas turbines has undergone steady development, leading to a reduction in emissions. In accordance with this design, a primary zone is operated with fuel-enriched combustion conditions, and the fuel/air mixture ratio is above the stoichiometric value. This permits limitation of the heat output and thus suppresses the production of thermal NOx. Downstream of the primary zone, a considerable amount of additional air is introduced into the combustion chamber. This takes place using exactly defined mixing geometries in order to convert the fuel/air mixture from rich to lean conditions. To prevent stoichiometric mixtures during the transition from rich to lean, which would lead to a significant temperature increase and consequently to a high thermal NOx production, the transition process is very fast and effective.

The design of the mixing zone of RQL combustion chambers is very important for achieving low NOx emissions. The state of the art shows designs in which the penetration of airflows from an inner and an outer annulus into the main burner flow is optimized to improve the mixing process and prevent high temperature peaks.

It is known from EP 0 676 590 B1 to provide air inlet recesses on the circumference in a row on the outer and the inner combustion chamber walls, with low-diameter air inlet openings being opposite high-diameter air inlet openings. The positioning of the air inlet openings is selected such that the airflows exiting the large air inlet recesses are aligned against a swirl direction of the air flowing out of the burner. This is intended to ensure effective mixing of air and fuel.

U.S. Pat. No. 6,260,359 B1 describes an embodiment, in which a second row of air inlet recesses is provided, the sizes of which differ in the circumferential direction.

A further design is shown in U.S. Pat. No. 6,675587 B2. Here too, the air guidance is selected such that it is aligned against the swirl direction of the flow exiting the burner, in order to improve the mixing of the air jets with the fuel/air mixture and to achieve a more even temperature distribution.

U.S. Pat. No. 7,363,763 B2 shows first and second groups of air inlet recesses, where the number of recesses in the groups differs and the spacings of the recesses in the circumferential direction vary.

U.S. Pat. No. 8,056,342 B2 describes a single-row mixing hole pattern for a combustion chamber, in which three different hole sizes are proposed per burner sector.

FIGS. 2 and 3 show embodiments in accordance with the state of the art a schematic representation of a gas-turbine engine as a whole being described with reference to FIG. 1.

FIG. 1 shows a schematic representation of a gas-turbine engine in accordance with the present invention The gas-turbine engine 110 in accordance with FIG. 1 is an example of a turbomachine where the invention can be used. The following however makes clear that the invention can also be used in other turbomachines. The engine 110 is of conventional design and includes in the flow direction, one behind the other, an air inlet 111, a fan 112 rotating inside a casing, an intermediate-pressure compressor 113, a high-pressure compressor 114, combustion chambers 115, a high-pressure turbine 116, an intermediate-pressure turbine 117 and a low-pressure turbine 118 as well as an exhaust nozzle 119, all of which being arranged about a center engine axis 101.

The intermediate-pressure compressor 113 and the high-pressure compressor 114 each include several stages, of which each has an arrangement extending in the circumferential direction of fixed and stationary guide vanes 120, generally referred to as stator vanes and projecting radially inwards from the engine casing 121 in an annular flow duct through the compressors 113, 114. The compressors furthermore have an arrangement of compressor rotor blades 122 which project radially outwards from a rotatable drum or disk 126 linked to hubs 127 of the high-pressure turbine 116 or the intermediate-pressure turbine 117, respectively.

The turbine sections 116, 117, 118 have similar stages, including an arrangement of fixed stator vanes 123 projecting radially inwards from the casing 121 into the annular flow duct through the turbines 116, 117, 118. and a subsequent arrangement of turbine blades 124 projecting outwards from a rotatable hub 127. The compressor drum or compressor disk 126 and the blades 122 arranged thereon, as well as the turbine rotor hub 127 and the turbine rotor blades 124 arranged thereon rotate about the engine axis 101 during operation.

FIG. 2 shows a simplified longitudinal sectional view of a burner 1 known from the state of the art including a fuel injection device 2. The arrow 3 shows the direction of the airflow through the fuel injection device 2 The combustion chamber includes a substantially annular inner combustion chamber wall 9 and a substantially annular outer combustion chamber wall 10. A burner flow 6, the course of which in an axial section is shown schematically in FIG. 2, flows into the combustion chamber formed by the combustion chamber walls 9 and 10. Air inlet recesses 11, 12, 13 are provided in the outer combustion chamber wall 10, while air inlet recesses 18, 19 and 20 are provided in the inner combustion chamber wall 9. Through these recesses pass, as shown in FIG. 2, airflows (mixed flows) 4 and 7 in a primary row of air inlet recesses and airflows 5 and 8 in a secondary row of air inlet recesses.

FIG. 3 shows a sectional view in a plane parallel to an injection plane (perpendicular to the gas-turbine center axis). The reference numeral 17 shows an aerodynamic swirl formed by the fuel injection device 2. Perpendicularly to the combustion chamber walls 9, 10, mixing flows 14, 15, 16 exit from the outer combustion chamber wall and mixing flows 21, 22, 23 from the inner combustion chamber wall, bringing about a reduction in the swirl intensity of the flow. The air inlet recesses 11, 12, 13, 18, 19, 20 can be designed here with differing diameters, as shown by the top view in the lower half of FIG. 3, so that differing air volume flows enter. The reference numerals 24 and 25 each show sector limits of the combustion chamber sectors, which are each assigned in the circumferential direction to a fuel injection device 2 (fuel injector).

The value Sx1_a indicates the axial spacing of the air inlet recesses between a front wall of the fuel injection device and the position of a first array of air inlet recesses.

With the solutions known from the state of the art, it has proved difficult to achieve good and fast mixing from rich to lean combustion conditions on the basis of the discrete mixing concepts. The potential of NOx reduction in known RQL combustion chamber designs is therefore limited.

The object underlying the present invention is to provide a gas-turbine combustion chamber of the type specified at the beginning which, while being simply designed and easily and cost-effectively producible, features a high mixing effectiveness and a good transition from rich to lean combustion states with low pollutant emissions.

It is a particular object of the present invention to provide solution to the above problematics by a combination of the features of claim 1. Further advantageous embodiments of the present invention become apparent from the sub-claims.

In accordance with the invention, an advantageous mixing design of the transition zone for gas-turbine combustion chambers with low emission values is thus created. The solution in accordance with the invention is based on the theory that the aerodynamic swirl of the flow, which can be generated by specific configurations inside the fuel injection nozzles and by other swirl-generating components located upstream of the mixing zone, is used to improve the penetration by the air jets entering the combustion chamber, in order to achieve in this way an improvement of the mixing quality. In accordance with the invention, therefore, there is a direct assignment of the airflows entering via the air inlet recesses to the swirl flow caused by the fuel injection device and by other swirl-generating components located upstream of the mixing zone.

In accordance with the invention, it is thus provided that at least one of the respective air inlet recesses is designed for supplying air in the flow direction of the swirl. The supplied air thus supports the swirling motion of the air/fuel flow exiting the fuel injection device and other swirl-generating components located upstream of the mixing zone. This leads to improved and more effective mixing.

In a preferred development of the invention, the geometries of the air inlet recesses vary in the circumferential direction in accordance with the formation of the swirl flow, in order to support the latter and to improve mixing. It is furthermore advantageous in accordance with the invention to select the distribution of the sir inlet recesses in the circumferential direction such that a reinforcement of the swirl flow is brought about by the swirl effect of the fuel injection nozzles and of other swirl-generating components, so that the swirl flow course prevailing in the transition zone is improved.

In accordance with the invention, it is particularly favourable when the mutually assigned air inlet recesses of the outer combustion chamber wall and of the inner combustion chamber wall are each arranged radially to one another or are arranged offset from one another in the circumferential direction. It can furthermore be favourable in accordance with the invention when the individual air inlet recesses are arranged substantially in one radial plane or are located in radial planes axially offset from one another.

In accordance with the invention, the air inlet recesses can have identical or varying spacing in the circumferential direction, depending on the respective combustion chamber design.

The essential feature of the present invention is the provision of air inlet recesses with flow guidance walls (chutes) 31, where said flow guidance walls 31 are in accordance with the invention designed with or without an inclination relative to the gas-turbine center axis and/or to a radial plane and/or in the circumferential direction. The air inlet recesses are axially arranged in a single row, where the air inlet recesses can be arranged in differing axial stages or offsets. It is furthermore possible in accordance with the invention to provide differing numbers of air inlet recesses on the outer combustion chamber wall and on the inner combustion chamber wall, respectively.

The individual burners of the gas-turbine combustion chamber in accordance with the invention can each be designed with a swirl in the same or in the opposite direction. The invention can be combined with differing combustion chamber cooling concepts, for example effusion cooling, Z-ring cooling or a cooling embodiment provided with insulating tiles or panels.

Furthermore the invention is suitable for both rich and lean combustion concepts.

The present invention is described in the following in light of the accompanying drawing, showing exemplary embodiments. In the drawing,

FIG. 1 shows a schematic representation of a gas-turbine engine in accordance with the present invention,

FIG. 2 shows an axial sectional view of a gas-turbine combustion chamber in accordance with the state of the art,

FIG. 3 shows a partial radial sectional view or a partial top view, respectively, of the gas-turbine combustion chamber in accordance with the state of the art shown in FIG. 2,

FIG. 4 shows a representation of a first exemplary embodiment by analogy with FIG. 3,

FIG. 5 shows a representation of a second exemplary embodiment by analogy with FIG. 4,

FIG. 6 shows a representation of a third exemplary embodiment by analogy with FIG. 4,

FIG. 7 shows a representation of a further exemplary embodiment by analogy with FIG. 4,

FIG. 8 shows a representation of a further exemplary embodiment by analogy with FIG. 4,

FIG. 9 shows a representation of a further exemplary embodiment by analogy with FIG. 4,

FIG. 10 shows a representation of a further exemplary embodiment by analogy with FIG. 4,

FIG. 11 shows a representation of a further exemplary embodiment by analogy with FIG. 4, and

FIG. 12 shows an axial section of the first exemplary embodiment by analogy with FIG. 4.

In the description of the exemplary embodiments, identical parts are provided with the same reference numerals.

The lower halves of FIGS. 4 to 11 each show partial top views onto the outer combustion chamber wall in a schematic representation.

In the exemplary embodiment shown in FIG. 4, an arrangement with three air inlet recesses per sector is provided on the outer combustion chamber wall or per fuel injection device 2. The same applies for the exemplary embodiment in FIG. 5.

FIG. 4 shows a single-row arrangement of the respective air inlet recesses 11. 12, 13 and 18. 19, 20, respectively. The individual air inlet recesses have differing diameters, such that the largest air volume flows 16, 23 respectively are introduced through the air inlet recesses 13 and 20 tangentially or almost tangentially to the swirl flow 17, while the air inlet recesses 11 and 18 each have a lower diameter, so that the flows 14 and 21 have a lower mass flow. The airflows 15 and 22 entering through the air inlet recesses 12 and 19 have, due to a smaller diameter of the air inlet recesses 12 and 19, a lower volume flow too. It is advantageous for the air inlet recesses 11 and 12 to have the same diameter, and this also applies to the air inlet recesses 18 and 19.

FIG. 4 shows the spacing Sy1_a of the air inlet recesses 11 and 12 in the circumferential direction on the outer combustion chamber wall 10, and the spacing Sy2_a of the air inlet recesses 12 and 13 in the circumferential direction on the outer combustion chamber wall 10. The spacing Sy1_i shows the spacing of the air inlet recesses 18 and 19 in the circumferential direction on the inner combustion chamber wall 9, and the spacing Sy2_i shows the spacing of the air inlet recesses 19 and 20 in the circumferential direction on the inner combustion chamber wall 9, The spacings Sy1_a and Sy2_a can be identical or can also differ, the same applying also for the spacings Sy1_i and Sy2_i.

FIG. 5 shows a modified exemplary embodiment in which the air inlet recesses on the outer and inner combustion chamber walls 20, 9 are offset from one another in the circumferential direction. The relative spacing in the circumferential direction is given as Sy1_ai and can differ between the individual air inlet recesses.

FIG. 6 shows a modified exemplary embodiment in which the air inlet recesses are provided with smaller bore diameters (11, 12, 18, 19) without chutes. The principal feature is however that the air inlet recesses with the high volume flows are provided with chutes 31 to reinforce the combustion chamber swirl (13, 22).

FIG. 7 shows a further exemplary embodiment with four air inlet recesses per sector on the outer and inner combustion chamber wall respectively. The size of the air inlet recess is designed similarly to the exemplary embodiment of FIG. 4, so that the largest mass flows impact the swirl flow 17 tangentially or substantially tangentially.

FIG. 8 shows a further exemplary embodiment where three air inlet recesses per sector are present on the outer combustion chamber wall 10 and two each air inlet recesses on the inner combustion chamber wall 9.

FIG. 9 shows a further design variant of the exemplary embodiment in FIG. 4, where two air inlet recesses per sector are provided with chutes 31 on the outer combustion chamber wall and on the inner combustion chamber wall. respectively.

FIG. 10 shows a further design variant of the exemplary embodiments of FIGS. 4 to 9, where the air inlet recesses on the outer combustion chamber wall and on the inner combustion chamber wall, respectively as well as the chutes are designed with non-circular flow cross-sections, for example as oblong holes.

FIG. 11 shows a further design variant of the exemplary embodiments of FIGS. 4 to 10, where the chutes are designed in the circumferential direction for reinforcing the swirl effect.

FIG. 12 shows for the exemplary embodiment of FIG. 4 an axial section of the combustion chamber with straight combustion chamber wall in preferably effusion-cooled design, where the mixing flows of differing intensity are indicated.

LIST OF REFERENCE NUMERALS

-   1 Burner -   2 Fuel injection device -   3 Airflow through the fuel injection device -   4 Mixed flow through a primary row of the outer combustion chamber     wall -   5 Mixed flow through a secondary row of the outer combustion chamber     wall -   6 Burner flow -   7 Mixed flow through a primary row of the inner combustion chamber     wall -   8 Mixed flow through a secondary row of the inner combustion chamber     wall -   9 Inner combustion chamber wall -   10 Outer combustion chamber wall -   11 Left-hand air inlet recess of the outer combustion chamber wall     (three air inlet recesses per sector area) -   12 Middle air filet recess of the outer combustion chamber wall     (three air inlet recesses per sector area) -   12 a Additional middle air inlet recess of the outer combustion     chamber wall (four air inlet recesses per sector area) -   13 Right-hand air inlet recess of the outer combustion chamber wall     (three air inlet recesses per sector area) -   14 Mixed flow through the left-hand air inlet recess of the outer     combustion chamber wall (three air inlet recesses per sector area) -   15 Mixed flow through the middle air inlet recess of the outer     combustion chamber wall (three air inlet recesses per sector area) -   16 Mixed flow through the right-hand air inlet recess of the outer     combustion chamber wall (three air inlet recesses per sector area) -   17 Aerodynamic swirl generated by the fuel injection device -   18 Right-hand air inlet recess of the inner combustion chamber wall     (three air inlet recesses per sector area) -   19 Middle air inlet recess of the inner combustion chamber wall     (three air inlet recesses per sector area) -   19 a Additional middle air inlet recess of the inner combustion     chamber wall (four air inlet recesses per sector area) -   20 Left-hand air inlet recess of the inner combustion chamber wail     (three air inlet recesses per sector area) -   21 Mixed flow through the right-hand air inlet recess of the inner     combustion chamber wall (three air inlet recesses per sector area) -   22 Mixed flow through the middle air inlet recess of the inner     combustion chamber wall (three air inlet recesses per sector area) -   23 Mixed flow through the left-hand air inlet recess of the inner     combustion chamber wall (three air inlet recesses per sector area) -   24 Left-hand sector limit -   25 Right-hand sector limit -   26 Position of primary row of air inlet recesses of the outer     combustion chamber wall -   27 Position of secondary row of air inlet recesses of the outer     combustion chamber wall -   28 Left-hand air inlet recess of the outer combustion chamber wall     (three air inlet recesses per sector area, secondary row) -   29 Middle air inlet recess of the outer combustion chamber wall     (three air inlet recesses per sector area, secondary row) -   30 Right-hand air inlet recess of the outer combustion chamber     (three air inlet recesses per sector area, secondary row) -   31 Flow guidance wall (chute) -   101 Engine axis -   110 Gas-turbine engine -   111 Air inlet -   112 Fan rotating inside the casing -   113 Intermediate-pressure compressor -   114 High-pressure compressor -   115 Combustion chambers -   116 High-pressure turbine -   117 Intermediate-pressure turbine -   118 Low-pressure turbine -   119 Exhaust nozzle -   120 Guide vanes -   121 Engine casing -   122 Compressor rotor blades -   123 Stator vanes -   124 Turbine blades -   126 Compressor drum or disk -   127 Turbine rotor hub -   Sy1_a Spacing of the air inlet recesses in the circumferential     direction on the outer combustion chamber wall -   Sy2_a Spacing of further air inlet recesses in the circumferential     direction on the outer combustion chamber wall -   Sy1_i Spacing of the air inlet recesses in the circumferential     direction on the inner combustion chamber wall -   Sy1_i Spacing of further air inlet recesses in the circumferential     direction on the inner combustion chamber wall

Sy1_ai Relative spacing of the opposite air inlet recesses in the circumferential direction

S1_a Axial spacing of the air inlet recesses between a front wall of the fuel injection device and the position of an air inlet recess on the outer combustion chamber wall

Sx1_i Axial spacing of the air inlet recesses between a front wall of the fuel injection device and the position of an air inlet recess on the inner combustion chamber wall

Sx2_a Axial spacing of the air inlet recesses between a front wall of the fuel injection device and the position of a further air inlet recess on the outer combustion chamber wall

Sx2_i Axial spacing of the air inlet recesses between a front wall of the fuel injection device and the position of a further air inlet recess on the inner combustion chamber wall 

1. Gas-turbine combustion chamber with an outer combustion chamber wall concentric to a gas-turbine center axis and with an inner combustion chamber wall, with several burners arranged spread over the circumference of the combustion chamber, and with air inlet recesses which in at least one radial plane are provided spread over the circumference on an outer combustion chamber wall and on an inner combustion chamber wall, where the burner is designed to form a flow provided with a swirl and where air inlet recesses assigned to a burner are dimensioned in differing size in order to generate airflows of differing size, characterized in that at least one of the respective air inlet recesses is designed for supplying air in the flow direction of the swirl, where at least one each air inlet recess is provided with flow guidance walls such that on the outer and inner combustion chamber walls one each air inlet recess is provided with a large air mass flow and the other recesses are provided with a reduced, however identical, air mass flow and that the large air inlet recesses are provided with flow guidance walls to intensify the mixture by a high jet pulse, with the swirl having the same direction as the combustion chamber flow.
 2. Gas-turbine combustion chamber in accordance with claim 1, characterized in that the mutually assigned air inlet recesses of the outer combustion chamber wall and of the inner combustion chamber wall are each arranged radially to one another or are arranged offset from one another in the circumferential direction.
 3. Gas-turbine combustion chamber in accordance with claim 1, characterized in that the mutually assigned air inlet recesses of the outer combustion chamber wall and of the inner combustion chamber wall have identical or varying spacing in the circumferential direction.
 4. Gas-turbine combustion chamber in accordance with claim 1, characterized in that all air inlet recesses are provided with flow guidance walls (chutes), where said flow guidance walls are designed with or without an inclination of the flow direction relative to the gas-turbine center axis and/or to a radial plane and/or to the circumferential direction, and are of identical or differing geometrical design.
 5. Gas-turbine combustion chamber in accordance with claim 1, characterized in that the air inlet recesses are arranged in axial stages to one another or are offset from one another.
 6. Gas-turbine combustion chamber in accordance with claim 1, characterized in that differing numbers of air inlet recesses are provided on the outer combustion chamber wall and on the inner combustion chamber wall.
 7. Gas-turbine combustion chamber in accordance with claim 1, characterized in that the individual burners are each designed for the formation of a swirl in the same or in the opposite direction to adjacent burners.
 8. Gas-turbine combustion chamber in accordance with claim 1, characterized in that the burner is designed as a lean mix burner or as a rich mix burner.
 9. Gas-turbine combustion chamber in accordance with claim 1, characterized in that one or several air inlet recesses are provided on the inner and outer combustion chamber walls per burner sector, with the number of air inlet recesses on the inner and outer combustion chamber walls being identical or differing.
 10. Gas-turbine combustion chamber in accordance with claim 1, characterized in that the swirl of the burner flow is partially or entirely generated by swirlers outside the fuel injection nozzles and that at least one of the respective air inlet recesses is designed for supplying air in the flow direction of the swirl.
 11. Gas-turbine combustion chamber in accordance with claim 1, characterized in that one or more air inlet recesses on the inner and outer combustion chamber walls are provided with non-circular air inlet recesses, for example as oblong holes, where at least the respectively largest air inlet recesses are provided with flow guidance walls. 